Compressor rotor stack assembly for gas turbine engine

ABSTRACT

A compressor rotor assembly including a plurality of rotor disks axially spaced from each other, each rotor disk extending radially from an inner end to an outer end. Also included is a spacer extending axially from each rotor disk to engage an adjacent spacer extending from an adjacent rotor disk, the spacer and adjacent spacer disposed proximate the outer end of the respective rotor disks, the spacers forming an outer backbone of the compressor rotor assembly. Further included is an inner backbone of the compressor rotor assembly, the inner backbone comprising a plurality of backbone segments, each of the backbone segments extending axially from each rotor disk to engage an adjacent backbone segment extending from an adjacent rotor disk, the backbone segment and the adjacent backbone segment disposed proximate the inner end of the respective rotor disks.

BACKGROUND

Exemplary embodiments pertain to the art of gas turbine engines and,more particularly, to a compressor rotor stack assembly for gas turbineengines.

A gas turbine engine typically includes one or more rotor stacksassociated with one or more sections of the engine. A rotor stack mayinclude several longitudinally spaced apart blade-carrying disks ofsuccessive stages of the section. A stator structure may includecircumferential stages of vanes longitudinally interspersed with therotor disks. The rotor disks are secured to each other against relativerotation and the rotor stack is secured against rotation relative toother components on its common spool (e.g., the low and highspeed/pressure spools of the engine).

Numerous systems have been used to tie rotor disks together. Forexample, a bore tied configuration may be employed, but a massivecompressive load is put on the rotor backbone, which is locatedproximate a radially outer end of the rotor disks. The backbone maybecome bowed, such that the compressive load puts the backbone intobending, thus forcing the backbone structure to be very thick. Gasturbine engines are being required to spin faster than previouslyrequired, thus emphasizing the need to reduce rotor stress byeliminating as much mass as possible from high radius regions.

BRIEF DESCRIPTION

Disclosed is a compressor rotor assembly including a plurality of rotordisks axially spaced from each other, each rotor disk extending radiallyfrom an inner end to an outer end. Also included is a spacer extendingaxially from each rotor disk to engage an adjacent spacer extending froman adjacent rotor disk, the spacer and adjacent spacer disposedproximate the outer end of the respective rotor disks, the spacersforming an outer backbone of the compressor rotor assembly. Furtherincluded is an inner backbone of the compressor rotor assembly, theinner backbone comprising a plurality of backbone segments, each of thebackbone segments extending axially from each rotor disk to engage anadjacent backbone segment extending from an adjacent rotor disk, thebackbone segment and the adjacent backbone segment disposed proximatethe inner end of the respective rotor disks.

In addition to one or more of the features described above, or as analternative, further embodiments may include a tie shaft extendingaxially proximate the inner end of the plurality of rotor disks toaxially compress the plurality of rotor disks together.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the plurality of rotordisks are part of a high pressure compressor assembly.

In addition to one or more of the features described above, or as analternative, further embodiments may include that tie shaft is threadedto a compressor structure proximate a last stage of the plurality ofrotor disks.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the inner backbonedefines a substantially cylindrical member.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the plurality of rotordisks are not bolted together.

Also disclosed is a gas turbine engine including a compressor section, acombustion section, a turbine section and a rotor disk assembly of thecompressor section. The rotor disk assembly includes a plurality ofrotor disks, each extending radially from an inner end to an outer end.The rotor disk assembly also includes a spacer extending axially fromeach rotor disk to engage an adjacent spacer extending from an adjacentrotor disk, the spacer and adjacent spacer disposed proximate the outerend of the respective rotor disks, the spacers forming an outer backboneof the compressor rotor assembly. The rotor disk assembly furtherincludes an inner backbone of the compressor rotor assembly, the innerbackbone comprising a plurality of backbone segments, each of thebackbone segments extending axially from each rotor disk to engage anadjacent backbone segment extending from an adjacent rotor disk, thebackbone segment and the adjacent backbone segment disposed proximatethe inner end of the respective rotor disks. The rotor disk assembly yetfurther includes a tie shaft extending axially proximate the inner endof the plurality of rotor disks to axially compress the plurality ofrotor disks together.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the plurality of rotordisks are part of a high pressure compressor assembly.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the tie shaft isthreaded to a compressor structure proximate a last stage of theplurality of rotor disks.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the inner backbonedefines a substantially cylindrical member.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the plurality of rotordisks are not bolted together.

Further disclosed is a method of assembling a compressor rotor assembly.The method includes arranging a plurality of rotor disks in an axiallyspaced manner relative to each other. The method also includes engaginga spacer extending from each rotor disk with an adjacent spacerextending from an adjacent rotor disk, the spacers disposed proximate aradially outer end of the rotor disks. The method further includesengaging a backbone segment extending from each rotor disk with anadjacent backbone segment extending from an adjacent backbone segment toform an inner backbone, the backbone segments disposed proximate aradially inner end of the rotor disks. The method yet further includesaxially compressing the spacers and the inner backbone with a tie shaftwith a positive torque applied thereto.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the tie shaft isthreaded to a compressor structure proximate a last stage of theplurality of rotor disks.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the inner backbonedefines a substantially cylindrical member.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the plurality of rotordisks are not bolted together.

In addition to one or more of the features described above, or as analternative, further embodiments may include that the plurality of rotordisks are part of a high pressure compressor assembly.

BRIEF DESCRIPTION OF THE DRAWINGS

The following descriptions should not be considered limiting in any way.With reference to the accompanying drawings, like elements are numberedalike:

FIG. 1 is a partial cross-sectional view of a gas turbine engine; and

FIG. 2 is an elevational view of a rotor assembly of the gas turbineengine.

DETAILED DESCRIPTION

A detailed description of one or more embodiments of the disclosedapparatus and method are presented herein by way of exemplification andnot limitation with reference to the Figures.

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct, while the compressor section 24 drives air along a coreflow path C for compression and communication into the combustor section26 then expansion through the turbine section 28. Although depicted as atwo-spool turbofan gas turbine engine in the disclosed non-limitingembodiment, it should be understood that the concepts described hereinare not limited to use with two-spool turbofans as the teachings may beapplied to other types of turbine engines including three-spoolarchitectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through aspeed change mechanism, which in exemplary gas turbine engine 20 isillustrated as a geared architecture 48 to drive the fan 42 at a lowerspeed than the low speed spool 30. The high speed spool 32 includes anouter shaft 50 that interconnects a high pressure compressor 52 and highpressure turbine 54. A combustor 56 is arranged in exemplary gas turbine20 between the high pressure compressor 52 and the high pressure turbine54. An engine static structure 36 is arranged generally between the highpressure turbine 54 and the low pressure turbine 46. The engine staticstructure 36 further supports bearing systems 38 in the turbine section28. The inner shaft 40 and the outer shaft 50 are concentric and rotatevia bearing systems 38 about the engine central longitudinal axis Awhich is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion. It will be appreciated that each of the positions of the fansection 22, compressor section 24, combustor section 26, turbine section28, and fan drive gear system 48 may be varied. For example, gear system48 may be located aft of combustor section 26 or even aft of turbinesection 28, and fan section 22 may be positioned forward or aft of thelocation of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five (5:1). Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present disclosure isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and35,000 feet (10,688 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5 m/sec).

Referring now to FIG. 2, a portion of the compressor section 24 isillustrated and may be referred to as a compressor rotor assembly or acompressor rotor stack assembly. The compressor rotor assembly 24includes a plurality of axially spaced rotor disks 60. Each illustratedrotor disk is one of a plurality of circumferentially spaced disks thatform various stages of the compressor section 24. It is to beappreciated that more or less stages may be present when compared to theillustrated embodiment. In some embodiments, the compressor rotorassembly 24 described herein is the high pressure compressor 52.

The rotor disks 60 are operatively coupled to one another in the mannerdescribed herein. In particular, a radially outer backbone 62 and aradially inner backbone 64 are axially compressed with at least onefastener. The radially outer backbone 62 is formed with a plurality ofspacers 68. The spacers 68 are integrally formed with, and extendaxially from the rotor disks 60. Specifically, each rotor disk 60includes a spacer extending therefrom that is engaged with an adjacentspacer that extends axially from an adjacent rotor disk. Similarly, theradially inner backbone 64 is formed of a plurality of backbone segments70 that are integrally formed with, and extend axially from the rotordisks 60. Specifically, each rotor disk includes a backbone segmentextending therefrom that is engaged with an adjacent backbone segmentthat extends axially from an adjacent rotor disk. As illustrated, thespacers 68 are located proximate a radially outer end 76 of the rotordisks 60 and the backbone segments 70 are located proximate a radiallyinner end 78 of the rotor disks 60.

As shown in the sectional view, the radially outer backbone 62 has anirregular geometric configuration, while the radially inner backbone 64is a cylindrical structure. A fastener, such as a tie shaft 72, isdisposed in engagement with the radially inner backbone 64 and theradially outer backbone 62 to axially compress the backbones in a mannerthat secures the rotor disks 60 to each other. By including the radiallyinner backbone 64, less of a compressive force must be fully absorbed bythe radially outer backbone 62. This allows the thickness of the spacers68 (i.e., radially outer backbone 62) to be significantly reduced,thereby reducing overall weight of the assembly. The radially innerbackbone 64 is substantially cylindrical so it may absorb thecompressive load applied by the tie shaft 72 in an efficient manner thatavoids bending of the backbone segments 70 (i.e., radially innerbackbone 64).

The tie shaft 72 is axially oriented along the radially inner end 78 ofthe rotor disks 60. In some embodiments, the tie shaft 72 is coupled tothe assembly by threading the tie shaft 72 to a portion of thecompressor structure, such as the portion referenced with numeral 80 inFIG. 2 and located proximate a last stage of the compressor section 24.In some embodiments, threaded engagement is also present proximate anupstream stage, such as with the portion of the compressor sectionreferenced with numeral 82. Additional or alternative engagementsections are contemplated.

Axially compressing the rotor disks with the tie shaft 72 alleviates theneed for bolted joints that are often required for rotor disk coupling.This significantly reduces the overall weight of the assembly.Therefore, the rotor disk coupling process does not include bolting thedisks together to achieve the weight reduction. Furthermore, no complexwelds are required, thereby simplifying the assembly process. Thecombination of an inner backbone to absorb the compressive load and thesecuring with a tie shaft 72 provides the advantages discussed above.

The term “about” is intended to include the degree of error associatedwith measurement of the particular quantity based upon the equipmentavailable at the time of filing the application. For example, “about”can include a range of ±8% or 5%, or 2% of a given value.

The terminology used herein is for the purpose of describing particularembodiments only and is not intended to be limiting of the presentdisclosure. As used herein, the singular forms “a”, “an” and “the” areintended to include the plural forms as well, unless the context clearlyindicates otherwise. It will be further understood that the terms“comprises” and/or “comprising,” when used in this specification,specify the presence of stated features, integers, steps, operations,elements, and/or components, but do not preclude the presence oraddition of one or more other features, integers, steps, operations,element components, and/or groups thereof.

While the present disclosure has been described with reference to anexemplary embodiment or embodiments, it will be understood by thoseskilled in the art that various changes may be made and equivalents maybe substituted for elements thereof without departing from the scope ofthe present disclosure. In addition, many modifications may be made toadapt a particular situation or material to the teachings of the presentdisclosure without departing from the essential scope thereof.Therefore, it is intended that the present disclosure not be limited tothe particular embodiment disclosed as the best mode contemplated forcarrying out this present disclosure, but that the present disclosurewill include all embodiments falling within the scope of the claims.

What is claimed is:
 1. A gas turbine engine comprising: a compressorsection; a combustion section; a turbine section; and a rotor diskassembly of the compressor section comprising: a plurality of rotordisks not bolted together, each extending radially from an inner end toan outer end; a spacer extending axially from each rotor disk to engagean adjacent spacer extending from an adjacent rotor disk, the spacer andadjacent spacer disposed proximate the outer end of a respective rotordisk, the spacers forming an outer backbone of the compressor rotorassembly; an inner backbone of the compressor rotor assembly, the innerbackbone comprising a plurality of backbone segments being axiallyaligned with one another and engaged with one another to define theinner backbone, each of the backbone segments extending axially fromeach rotor disk to engage an adjacent backbone segment extending from anadjacent rotor disk, each backbone segment of the plurality of backbonesegments being disposed proximate the inner end of the respective rotordisk; and a tie shaft extending axially proximate the inner end of theplurality of rotor disks to axially compress the plurality of rotordisks together.
 2. The gas turbine engine of claim 1, wherein theplurality of rotor disks are part of a high pressure compressorassembly.
 3. The gas turbine engine of claim 1, the tie shaft threadedto a compressor structure proximate a last stage of the plurality ofrotor disks.
 4. The gas turbine engine of claim 1, wherein the innerbackbone defines a substantially cylindrical member.
 5. A method ofassembling a compressor rotor assembly comprising: arranging a pluralityof rotor disks not bolted together in an axially spaced manner relativeto each other; engaging a spacer extending from each rotor disk with anadjacent spacer extending from an adjacent rotor disk, the spacersdisposed proximate a radially outer end of the rotor disks; engaging abackbone segment extending from each rotor disk with an adjacentbackbone segment extending from an adjacent rotor disk to form an innerbackbone, the backbone segments disposed proximate a radially inner endof the rotor disks, the backbone segments being axially aligned with oneanother and engaged with one another; and axially compressing thespacers and the inner backbone with a tie shaft with a positive torqueapplied thereto.
 6. The method of claim 5, wherein the tie shaft isthreaded to a compressor structure proximate a last stage of theplurality of rotor disks.
 7. The method of claim 5, wherein the innerbackbone defines a substantially cylindrical member.
 8. The method ofclaim 5, wherein the plurality of rotor disks are part of a highpressure compressor assembly.